Gas turbine engine with high speed low pressure turbine section and bearing support features

ABSTRACT

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan, a compressor section, and a turbine section including a fan drive turbine and a second turbine. The fan drive turbine has a first exit area at a first exit point and is rotatable at a first speed. A mid-turbine frame is positioned intermediate the fan drive turbine and the second turbine, and can include a bearing support. The second turbine has a second exit area at a second exit point and is rotatable at a second speed. A first performance quantity is defined as the product of the first speed squared and the first area. A second performance quantity is defined as the product of the second speed squared and the second area.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation-in-part of U.S. patent applicationSer. No. 13/446,510 filed Apr. 13, 2012, which claims priority to U.S.Provisional Application No. 61/619,124, filed Apr. 2, 2012, and is acontinuation-in-part of U.S. patent application Ser. No. 13/363,154,filed on Jan. 31, 2012 and entitled “Gas Turbine Engine With High SpeedLow Pressure Turbine Section.”

BACKGROUND OF THE INVENTION

This application relates to a gas turbine engine wherein the lowpressure turbine section is rotating at a higher speed and centrifugalpull stress relative to the high pressure turbine section speed andcentrifugal pull stress than prior art engines.

Gas turbine engines are known, and typically include a fan deliveringair into a low pressure compressor section. The air is compressed in thelow pressure compressor section, and passed into a high pressurecompressor section. From the high pressure compressor section the air isintroduced into a combustion section where it is mixed with fuel andignited. Products of this combustion pass downstream over a highpressure turbine section, and then a low pressure turbine section.

Traditionally, on many prior art engines the low pressure turbinesection has driven both the low pressure compressor section and a fandirectly. As fuel consumption improves with larger fan diametersrelative to core diameters it has been the trend in the industry toincrease fan diameters. However, as the fan diameter is increased, highfan blade tip speeds may result in a decrease in efficiency due tocompressibility effects. Accordingly, the fan speed, and thus the speedof the low pressure compressor section and low pressure turbine section(both of which historically have been coupled to the fan via the lowpressure spool), have been a design constraint. More recently, gearreductions have been proposed between the low pressure spool (lowpressure compressor section and low pressure turbine section) and thefan.

SUMMARY

A gas turbine engine according to an example of the present disclosureincludes a fan that has fewer than 26 fan blades, wherein a fan pressureratio across the fan blades is less than 1.45, measured across the fanblades alone, a compressor section that has a first compressor and asecond compressor, a bypass ratio greater than 10, a turbine sectionthat has a fan drive turbine and a second turbine, and a gear systemwith a gear reduction. The fan drive turbine drives the fan through thegear system, and a gear ratio of the gear reduction being greater than2.5. A mid-turbine frame is positioned intermediate the fan driveturbine and the second turbine, and has a first bearing supporting afirst shaft rotatable with the fan drive turbine in an overhung manner.The fan drive turbine has a first exit area at a first exit point and isrotatable at a first speed. The second turbine has a second exit area ata second exit point and is rotatable at a second speed. A firstperformance quantity is defined as the product of the first speedsquared and the first area. A second performance quantity is defined asthe product of the second speed squared and the second area. Aperformance ratio of the first performance quantity to the secondperformance quantity is between 0.5 and 1.5.

In a further embodiment of any of the foregoing embodiments, the fandrive turbine is a 3-stage to 6-stage turbine, and the second turbine isa 2-stage turbine.

In a further embodiment of any of the foregoing embodiments, the firstcompressor has 3 stages.

A further embodiment of any of the foregoing embodiments include a fandrive shaft interconnecting the gear system and the fan, and a framesupporting at least a portion of the fan drive shaft. The frame definesa frame transverse stiffness. A flexible support at least partiallysupports the gear system. The flexible support defines a supporttransverse stiffness with respect to the frame transverse stiffness. Thesupport transverse stiffness is less than about 50% of the frametransverse stiffness.

In a further embodiment of any of the foregoing embodiments, themid-turbine frame includes a guide vane positioned intermediate the fandrive turbine and the second turbine.

In a further embodiment of any of the foregoing embodiments, theperformance ratio is equal to or greater than 0.8.

In a further embodiment of any of the foregoing embodiments, the fandrive turbine includes an inlet, an outlet, and a fan drive turbinepressure ratio greater than 5, wherein the fan drive turbine pressureratio is a ratio of a pressure measured prior to the inlet as related toa pressure at the outlet prior to any exhaust nozzle.

In a further embodiment of any of the foregoing embodiments, themid-turbine frame includes a plurality of airfoils in a core airflowpath.

In a further embodiment of any of the foregoing embodiments, theperformance ratio is greater than or equal to 1.0.

In a further embodiment of any of the foregoing embodiments, the fanblades have a fan tip speed of less than 1150 ft/second, and the gearsystem is a planetary gear system.

In a further embodiment of any of the foregoing embodiments, theperformance ratio is less than or equal to 1.075.

In a further embodiment of any of the foregoing embodiments, the secondspeed is greater than twice the first speed.

In a further embodiment of any of the foregoing embodiments, themid-turbine frame has a second bearing supporting an outer periphery ofa second shaft rotatable with the second turbine.

In a further embodiment of any of the foregoing embodiments, the gearsystem is intermediate the fan drive turbine and the first compressorsuch that the fan and the first compressor are rotatable at a commonspeed.

In a further embodiment of any of the foregoing embodiments, the gearsystem is straddle-mounted by bearings.

A further embodiment of any of the foregoing embodiments include a fandrive shaft interconnecting the gear system and the fan, and a framesupporting at least a portion of the fan drive shaft. The frame definesa frame transverse stiffness and a frame lateral stiffness. A flexiblesupport at least partially supports the gear system. The flexiblesupport defines a support transverse stiffness with respect to the frametransverse stiffness and a support lateral stiffness with respect to theframe lateral stiffness. The support transverse stiffness is less thanabout 80% of the frame transverse stiffness. The support lateralstiffness is less than about 80% of the frame lateral stiffness.

In a further embodiment of any of the foregoing embodiments, the supporttransverse stiffness is less than about 50% of the frame transversestiffness, and the support lateral stiffness is less than about 50% ofthe frame lateral stiffness

A gas turbine engine according to an example of the present disclosureincludes a fan that has fewer than 26 fan blades, wherein a fan pressureratio across the fan blades is less than 1.45, measured across the fanblades alone, a turbine section that has a fan drive turbine and asecond turbine, and a gear system with a gear reduction. The fan driveturbine drives the fan through the gear system, and a gear ratio of thegear reduction is greater than 2.5. A mid-turbine frame is positionedintermediate the fan drive turbine and the second turbine. The fan driveturbine has a first exit area at a first exit point and is rotatable ata first speed. The second turbine has a second exit area at a secondexit point and is rotatable at a second speed, and a first performancequantity is defined as the product of the first speed squared and thefirst area. A second performance quantity is defined as the product ofthe second speed squared and the second area, and a performance ratio ofthe first performance quantity to the second performance quantity isbetween 0.8 and 1.5.

In a further embodiment of any of the foregoing embodiments, themid-turbine frame includes a guide vane positioned intermediate the fandrive turbine and the second turbine.

In a further embodiment of any of the foregoing embodiments, the fandrive turbine and second turbine are rotatable in opposed directions,and the guide vane is an air turning guide vane.

In a further embodiment of any of the foregoing embodiments, themid-turbine frame has a first bearing supporting a first shaft rotatablewith the fan drive turbine in an overhung manner.

In a further embodiment of any of the foregoing embodiments, the fandrive turbine is a 3-stage to 6-stage turbine, and the second turbine isa 2-stage turbine.

A further embodiment of any of the foregoing embodiments include abypass ratio greater than 10.

In a further embodiment of any of the foregoing embodiments, the fandrive turbine includes an inlet, an outlet, a fan drive turbine pressureratio greater than 5, the fan drive turbine pressure ratio being a ratioof a pressure measured prior to the inlet as related to a pressure atthe outlet prior to any exhaust nozzle. The first speed is greater than10,000 RPM, and the second speed is greater than 20,000 RPM.

In a further embodiment of any of the foregoing embodiments, the fandrive turbine includes an inlet, an outlet, and a fan drive turbinepressure ratio greater than 5, wherein the fan drive turbine pressureratio is a ratio of a pressure measured prior to the inlet as related toa pressure at the outlet prior to any exhaust nozzle. The fan driveturbine is a 3-stage to 6-stage turbine, and the second turbine is a2-stage turbine.

In a further embodiment of any of the foregoing embodiments, a bypassratio is greater than 10.

In a further embodiment of any of the foregoing embodiments, the fandrive turbine is a 3-stage turbine.

A further embodiment of any of the foregoing embodiments a fan driveshaft interconnecting the gear system and the fan, and a framesupporting at least a portion of the fan drive shaft. The frame definesa frame transverse stiffness, a flexible support at least partiallysupports the gear system. The flexible support defines a supporttransverse stiffness with respect to the frame transverse stiffness. Thesupport transverse stiffness is less than about 50% of the frametransverse stiffness.

In a further embodiment of any of the foregoing embodiments, theperformance ratio is greater than or equal to 1.0.

In a further embodiment of any of the foregoing embodiments, the secondspeed is greater than twice the first speed.

In a featured embodiment, a turbine section of a gas turbine engine hasa fan drive and second turbine sections. The fan drive turbine sectionhas a first exit area at a first exit point and is configured to rotateat a first speed. The second turbine section has a second exit area at asecond exit point and rotates at a second speed, which is faster thanthe first speed. A first performance quantity is defined as the productof the first speed squared and the first area. A second performancequantity is defined as the product of the second speed squared and thesecond area. A ratio of the first performance quantity to the secondperformance quantity is between about 0.5 and about 1.5. A mid-turbineframe is positioned intermediate the fan drive and second turbinesections, and has a first bearing supporting an outer periphery of afirst shaft rotating with the second turbine section.

In another embodiment according to the previous embodiment, themid-turbine frame also includes a second bearing supporting an outerperiphery of a second shaft rotating with the fan drive turbine section.The second bearing supports an intermediate portion of the second spool.

In another embodiment according to any of the previous embodiments, theratio is above or equal to about 0.8.

In another embodiment according to any of the previous embodiments, thefan drive turbine section has at least 3 stages.

In another embodiment according to any of the previous embodiments, thefan drive turbine section has up to 6 stages.

In another embodiment according to any of the previous embodiments, thesecond turbine section has 2 or fewer stages.

In another embodiment according to any of the previous embodiments, apressure ratio across the fan drive turbine section is greater thanabout 5:1.

In another embodiment according to any of the previous embodiments, themid-turbine frame is provided with a guide vane positioned intermediatethe fan drive and second turbine sections.

In another embodiment according to any of the previous embodiments, thefan drive and second turbine sections will rotate in opposed directions.The guide vane is a turning guide vane.

In another featured embodiment, a gas turbine engine has a fan, acompressor section in fluid communication with the fan, a combustionsection in fluid communication with the compressor section, and aturbine section in fluid communication with the combustion section. Theturbine section includes a fan drive turbine section and a secondturbine section. The fan drive turbine section has a first exit area ata first exit point and is configured to rotate at a first speed. Thesecond turbine section has a second exit area at a second exit point androtates at a second speed, which is higher than the first speed. A firstperformance quantity is defined as the product of the first speedsquared and the first area. A second performance quantity is defined asthe product of the second speed squared and the second area. A ratio ofthe first performance quantity to the second performance quantity isbetween about 0.5 and about 1.5. The second turbine section is supportedby a first bearing in a mid-turbine frame.

In another embodiment according to the previous embodiment, the ratio isabove or equal to about 0.8.

In another embodiment according to any of the previous embodiments, thecompressor section includes first and second compressor sections. Thefan drive turbine section and the first compressor section will rotatein a first direction. The second turbine section and the secondcompressor section will rotate in a second opposed direction.

In another embodiment according to any of the previous embodiments, agear reduction is included between the fan and a shaft driven by the fandrive turbine section such that the fan will rotate at a lower speedthan the fan drive turbine section.

In another embodiment according to any of the previous embodiments, thesecond turbine section and second compressor section arestraddle-mounted by bearings supported on an outer periphery of a shaftrotating with the second compressor section and the second turbinesection.

In another embodiment according to any of the previous embodiments, themid-turbine frame further includes a second bearing supporting an outerperiphery of a shaft rotating with the fan drive turbine section.

In another embodiment according to any of the previous embodiments, thesecond bearing supports an intermediate portion of a shaft that willrotate with the fan drive turbine section and the first compressorsection.

In another featured embodiment, a gas turbine engine has a fan, acompressor section in fluid communication with the fan, a combustionsection in fluid communication with the compressor section, and aturbine section in fluid communication with the combustion section. Theturbine section includes fan drive and second turbine sections. The fandrive turbine section has a first exit area at a first exit point and isconfigured to rotate at a first speed. The second turbine section has asecond exit area at a second exit point and rotates at a second speed,which is higher than the first speed. A first performance quantity isdefined as the product of the first speed squared and the first area. Asecond performance quantity is defined as the product of the secondspeed squared and the second area. A ratio of the first performancequantity to the second performance quantity is between about 0.5 andabout 1.5. The compressor section includes first and second compressorsections, where the fan drive turbine section and the first compressorsection will rotate in a first direction and the second turbine sectionand the second compressor section will rotate in a second opposeddirection. A gear reduction is included between the fan and the firstcompressor section, such that the fan will rotate at a lower speed thanthe fan drive turbine section, and in the second opposed direction. Agear ratio of the gear reduction is greater than about 2.3.

In another embodiment according to the previous embodiment, the ratio isabove or equal to about 0.8.

In another embodiment according to any of the previous embodiments, amid-turbine frame is positioned intermediate the fan drive and secondturbine sections. The mid-turbine frame has a first bearing supportingan outer periphery of a first shaft rotating with the second turbinesection.

In another embodiment according to any of the previous embodiments, thefirst shaft is supported on a second bearing on its outer periphery,with the second bearing mounted to static structure.

These and other features of this disclosure will be better understoodupon reading the following specification and drawings, the following ofwhich is a brief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 shows a gas turbine engine.

FIG. 2 schematically shows the arrangement of the low and high spool,along with the fan drive.

FIG. 3 shows a schematic view of a mount arrangement for an engine suchas shown in FIGS. 1 and 2.

FIG. 4 shows another embodiment of a gas turbine engine.

FIG. 5 shows yet another embodiment of a gas turbine engine.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-turbine turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B whilethe compressor section 24 drives air along a core flow path C forcompression and communication into the combustor section 26 thenexpansion through the turbine section 28. Although depicted as aturbofan gas turbine engine in the disclosed non-limiting embodiment, itshould be understood that the concepts described herein are not limitedto use with turbofans as the teachings may be applied to other types ofturbine engines including three-turbine architectures.

The engine 20 generally includes a low speed spool 30 and a high speedspool 32 mounted for rotation about an engine central longitudinal axisA relative to an engine static structure 36 via several bearing systems38. It should be understood that various bearing systems 38 at variouslocations may alternatively or additionally be provided.

The low speed spool 30 generally includes an innermost shaft 40 thatinterconnects a fan 42, a low pressure (or first) compressor section 44and a low pressure (or first) turbine section 46. Note, turbine section46 will also be called a fan drive turbine section. The inner shaft 40is connected to the fan 42 through a geared architecture 48 to drive thefan 42 at a lower speed than the fan drive turbine 46. The high speedspool 32 includes a more outer shaft 50 that interconnects a highpressure (or second) compressor section 52 and high pressure (or second)turbine section 54. A combustor 56 is arranged between the high pressurecompressor section 52 and the high pressure turbine section 54. Amid-turbine frame 57 of the engine static structure 36 is arrangedgenerally between the high pressure turbine section 54 and the lowpressure turbine section 46. The mid-turbine frame 57 further supportsbearing systems 38 in the turbine section 28. As used herein, the highpressure turbine section experiences higher pressures than the lowpressure turbine section. A low pressure turbine section is a sectionthat powers a fan 42. The inner shaft 40 and the outer shaft 50 areconcentric and rotate via bearing systems 38 about the engine centrallongitudinal axis A which is collinear with their longitudinal axes. Thehigh and low spools can be either co-rotating or counter-rotating.

The core airflow C is compressed by the low pressure compressor section44 then the high pressure compressor section 52, mixed and burned withfuel in the combustor 56, then expanded over the high pressure turbinesection 54 and low pressure turbine section 46. The mid-turbine frame 57includes airfoils 59 (one shown in FIG. 1) which are in the core airflowpath. The turbine sections 46, 54 rotationally drive the respective lowspeed spool 30 and high speed spool 32 in response to the expansion.

The engine 20 in one example is a high-bypass geared aircraft engine.The bypass ratio is the amount of air delivered into bypass path Bdivided by the amount of air into core path C. In a further example, theengine 20 bypass ratio is greater than about six (6), with an exampleembodiment being greater than ten (10), the geared architecture 48 is anepicyclic gear train, such as a planetary gear system or other gearsystem, with a gear reduction ratio of greater than about 2.3 and thelow pressure turbine section 46 has a pressure ratio that is greaterthan about 5. In some embodiments, the bypass ratio is less than orequal to about 22.0, and the gear reduction is less than or equal toabout 4.5. In one disclosed embodiment, the engine 20 bypass ratio isgreater than about ten (10:1), the fan diameter is significantly largerthan that of the low pressure compressor section 44, and the lowpressure turbine section 46 has a pressure ratio that is greater thanabout 5:1. In some embodiments, the low pressure turbine section 46 hasa pressure ratio that is less than or equal to about 30. In someembodiments, the high pressure turbine section may have two or fewerstages. In contrast, the low pressure turbine section 46, in someembodiments, has between 3 and 6 stages. Further the low pressureturbine section 46 pressure ratio is total pressure measured prior toinlet of low pressure turbine section 46 as related to the totalpressure at the outlet of the low pressure turbine section 46 prior toan exhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.5:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (“TSFC”). TSFC is the industry standardparameter of the rate of lbm of fuel being burned per hour divided bylbf of thrust the engine produces at that flight condition. “Low fanpressure ratio” is the ratio of total pressure across the fan bladealone, before the fan exit guide vanes. The low fan pressure ratio asdisclosed herein according to one non-limiting embodiment is less thanabout 1.45, and is greater than or equal to about 1.1. “Low correctedfan tip speed” is the actual fan tip speed in ft/sec divided by anindustry standard temperature correction of [(Ram Air Temperature degR)/518.7)^0.5]. The “Low corrected fan tip speed” as disclosed hereinaccording to one non-limiting embodiment is less than about 1150ft/second, and is greater than or equal to about 850 ft/second. Further,the fan 42 may have 26 or fewer blades.

An exit area 400 is shown, in FIG. 1 and FIG. 2, at the exit locationfor the high pressure turbine section 54 is the annular area of the lastblade of turbine section 54. An exit area for the low pressure turbinesection is defined at exit 401 for the low pressure turbine section andis the annular area defined by the blade of that turbine section 46. Asshown in FIG. 2, the turbine engine 20 may be counter-rotating. Thismeans that the low pressure turbine section 46 and low pressurecompressor section 44 rotate in one direction (“−”), while the highpressure spool 32, including high pressure turbine section 54 and highpressure compressor section 52 rotate in an opposed (“+”) direction. Thegear reduction 48, which may be, for example, an epicyclic transmission(e.g., with a sun, ring, and planet gears), is selected such that thefan 42 rotates in the same direction (“+”) as the high spool 32. Withthis arrangement, and with the other structure as set forth above,including the various quantities and operational ranges, a very highspeed can be provided to the low pressure spool. Low pressure turbinesection and high pressure turbine section operation are often evaluatedlooking at a performance quantity which is the exit area for the turbinesection multiplied by its respective speed squared. This performancequantity (“PQ”) is defined as:PQ_(ltp)=(A _(lpt) ×V _(lpt) ²)  Equation 1:PQ_(hpt)=(A _(hpt) ×V _(hpt) ²)  Equation 2:where A_(lpt) is the area of the low pressure turbine section at theexit thereof (e.g., at 401), where V_(lpt) is the speed of the lowpressure turbine section, where A_(hpt) is the area of the high pressureturbine section at the exit thereof (e.g., at 400), and where V_(hpt) isthe speed of the low pressure turbine section. As known, one wouldevaluate this performance quantity at the redline speed for each turbinesection.

Thus, a ratio of the performance quantity for the low pressure turbinesection compared to the performance quantify for the high pressureturbine section is:(A _(lpt) ×V _(lpt) ²)/(A _(hpt) ×V _(hpt)²)=PQ_(ltp)/PQ_(hpt)  Equation 3:In one turbine embodiment made according to the above design, the areasof the low and high pressure turbine sections are 557.9 in² and 90.67in², respectively. Further, the redline speeds of the low and highpressure turbine sections are 10179 rpm and 24346 rpm, respectively,such that the speed of the high pressure turbine section is more thantwice the speed of the low pressure section, and such that the speeds ofthe low and high pressure turbine sections being greater than 10000 and20000 rpm, respective. That is, the high speed is more than twice thelow speed and less than 2.8 times the low speed. Thus, using Equations 1and 2 above, the performance quantities for the low and high pressureturbine sections are:PQ_(ltp)=(A _(lpt) ×V _(lpt) ²)=(557.9 in²)(10179 rpm)²=57805157673.9in² rpm²  Equation 1:PQ_(hpt)=(A _(hpt) ×V _(hpt) ²)=(90.67 in²)(24346 rpm)²=53742622009.72in² rpm²  Equation 2:and using Equation 3 above, the ratio for the low pressure turbinesection to the high pressure turbine section is:Ratio=PQ_(ltp)/PQ_(hpt)=57805157673.9 in² rpm²/53742622009.72 in²rpm²=1.075

In another embodiment, the ratio was about 0.5 and in another embodimentthe ratio was about 1.5. With PQ_(ltp)/PQ_(hpt) ratios in the 0.5 to 1.5range, a very efficient overall gas turbine engine is achieved. Morenarrowly, PQ_(ltp)/PQ_(hpt) ratios of above or equal to about 0.8 aremore efficient. Even more narrowly, PQ_(ltp)/PQ_(hpt) ratios above orequal to 1.0 are even more efficient. As a result of thesePQ_(ltp)/PQ_(hpt) ratios, in particular, the turbine section can be mademuch smaller than in the prior art, both in diameter and axial length.In addition, the efficiency of the overall engine is greatly increased.

The low pressure compressor section is also improved with thisarrangement, and behaves more like a high pressure compressor sectionthan a traditional low pressure compressor section. It is more efficientthan the prior art, and can provide more compression in fewer stages.The low pressure compressor section may be made smaller in radius andshorter in length while contributing more toward achieving the overallpressure ratio design target of the engine.

As shown in FIG. 3, the engine as shown in FIGS. 1 and 2 may be mountedsuch that the high pressure turbine 54 is supported on a rear end by amid-turbine frame 110. The mid-turbine frame 110 may be provided with aguide vane 112 that is an air turning vane. Since the high pressureturbine 54 and the low pressure or fan drive turbine 46 rotate inopposed directions, the use of the turning vane intermediate the twowill ensure that the gases leaving the high pressure turbine 54 approachthe low pressure turbine 46 traveling in the proper direction. As isclear from FIG. 3, the mid-turbine frame 110 also includes a bearing 116which supports a shaft that rotates with the low spool 30 in an“overhung” manner. That is, the bearing 116 is at an intermediateposition on the shaft, rather than adjacent the end.

Static structure 102 and 108 support other bearings 100 and 110 tosupport the shafts driven by spools 30 and 32 on the compressor end. Thehigh pressure turbine 54 can be said to be “straddle-mounted” due to thebearings 110 and 114 on the outer periphery of the shaft 32.

FIG. 4 shows an embodiment 200, having a gear reduction 248 intermediatea low pressure (or first) compressor section 244 and a shaft 230 drivenby a low pressure turbine section 246. Embodiment 200 can be utilizedand/or combined with the features of the engine 20 as shown in FIGS. 1to 3, for example, with like reference numerals designating likeelements where appropriate and reference numerals with the addition ofone-hundred or multiples thereof designating modified elements that areunderstood to incorporate the same features and benefits of thecorresponding original elements.

The gear reduction 248 includes a sun gear 248A attached to the shaft230. The sun gear 248A can be mounted to a flexible input 231 which isattached to the shaft 230. The gear reduction 248 can include a splineinterface 231A, in which flexible input 231 has a spline which mateswith and engages an inner periphery of the sun gear 248A. Accordingly,the sun gear 248A can be driven by the spline interface 231A of theflexible input 231. Surrounding the sun gear 248A is a plurality ofplanet gears 248B supported on bearings 248C attached to a carrier 248Dmounted to a fan drive shaft 243. The planet gears 248B are surroundedon a radially outward side by a ring gear 248E. The fan drive shaft 243interconnects an output 249 of the gear reduction 248 and the fan 242,with the fan 242 and the low pressure compressor section 244 beingdriven by the output 249 of the gear reduction 248. In the illustratedembodiment of FIG. 4, the output 249 is the carrier 248D of the gearreduction 248, with the carrier 248D being rotatable about the engineaxis A and the ring gear 248E being fixed. The low pressure compressor244 can be driven by the shaft 230 such that the low pressure compressor244 and the low pressure turbine 246 are rotatable at a common speed anda common direction. In another example, the fan 242 and the low pressurecompressor section 244 are coupled to the gear reduction 248 such thatthe fan 242 and the low pressure compressor section 244 are rotatable atdifferent speeds and/or in different directions.

The ring gear 248E can be attached to the engine static structure 236through a flexible support 251 which at least partially supports thegear reduction 248. The static structure 236 includes a bearing supportor frame 236A which supports at least a portion of the fan drive shaft243 via a fan shaft roller bearing 217 and a fan shaft thrust bearing218. The gear reduction 248 connects to the fan drive shaft 243 axiallyforward of the fan shaft roller bearing 217 and axially rearward of thefan shaft thrust bearing 218 in order to allow the gear reduction 248 tobe at least partially axially aligned with the low pressure compressor244. Alternatively, the fan shaft roller bearings 217 could be locatedaxially forward of the gear reduction 248 and the fan shaft thrustbearing 218 could be located axially aft of the gear reduction 248. Thebearings 217 and 218 are positioned on opposite sides of the gearreduction 248 relative to engine axis A and support the gear reduction248 in a “straddle-mounted” manner. In the illustrated embodiment ofFIG. 4, the fan shaft roller bearing 217 supports an aft portion of thecarrier 248D.

The frame 236A defines a frame lateral stiffness and a frame transversestiffness. It should be understood that the term “lateral” as definedherein is generally transverse to the engine axis A, and the term“transverse” refers to a pivotal bending movement with respect to theengine axis A which typically absorbs deflection applied to the gearreduction 248. The flexible input 231 and the flexible support 251 eachcan be arranged to define a respective support/input lateral stiffnessand a support/input transverse stiffness.

In examples, the support transverse stiffness and/or the inputtransverse stiffness are less than the frame transverse stiffness. Insome examples, the support lateral stiffness and/or the input lateralstiffness are less than the frame transverse stiffness. In one example,both the support lateral stiffness and the input lateral stiffness areless than about 80% of the frame lateral stiffness, or more narrowlyless than about 50%, with the lateral stiffness of the entire gearreduction 248 being controlled by this lateral stiffness relationship.Alternatively, or in addition to this relationship, both the supporttransverse stiffness and the input transverse stiffness are each lessthan about 80% of the frame transverse stiffness, or more narrowlybetween 80% and 50%, less than about 65%, or less than about 50%, withthe transverse stiffness of the entire gear reduction 248 beingcontrolled by this transverse stiffness relationship. In some examples,the support lateral stiffness and/or the input lateral stiffness areless than about 20% of the frame lateral stiffness. In other examples,the support transverse stiffness and/or the input transverse stiffnessare less than about 20% of the frame transverse stiffness.

FIG. 5 shows an embodiment 300 with gear reduction 348. In theillustrated example of FIG. 5, the output 349 is the ring gear 348E,with the fan drive shaft 343 mechanically attached to a fan rotor 342Aof the fan 342 such that the fan 342 and the low pressure compressorsection 344 are rotatable at a common speed and in a common direction.The carrier 348D is attached to the engine static structure 336 througha flexible support 351. The gear reduction 348 includes a flexibleoutput coupling 247 which interconnects output 249 of the gear reduction348 and the fan drive shaft 343. The flexible input 331, the flexibleoutput coupling 347, and the flexible support 351 work together tomaintain alignment of the gear reduction 348 and can facilitate thesegregation of vibrations and other transients between the variouscomponents during operation of the gas turbine engine 20.

While this invention has been disclosed with reference to oneembodiment, it should be understood that certain modifications wouldcome within the scope of this invention. For that reason, the followingclaims should be studied to determine the true scope and content of thisinvention.

What is claimed is:
 1. A gas turbine engine comprising: a fan havingfewer than 26 fan blades; a compressor section including a firstcompressor and a second compressor; a turbine section including a fandrive turbine and a second turbine; a gear system with a gear reduction,the fan drive turbine driving the fan through the gear system; amid-turbine frame positioned intermediate the fan drive turbine and thesecond turbine, and having a first bearing supporting a first shaftrotatable with the fan drive turbine in an overhung manner; wherein thefan drive turbine has a first exit area at a first exit point and isrotatable at a first speed, the second turbine has a second exit area ata second exit point and is rotatable at a second speed, said first andsecond speeds being redline speeds; and wherein a first performancequantity is defined as the product of the first speed squared and thefirst area, a second performance quantity is defined as the product ofthe second speed squared and the second area, and a performance ratio ofthe first performance quantity to the second performance quantity isbetween 0.5 and 1.5.
 2. The gas turbine engine as set forth in claim 1,wherein the fan drive turbine is a 3-stage to 6-stage turbine, and thesecond turbine is a 2-stage turbine.
 3. The gas turbine engine as setforth in claim 2, wherein the first compressor has 3 stages.
 4. The gasturbine engine as set forth in claim 2, wherein the mid-turbine frameincludes a guide vane positioned intermediate the fan drive turbine andthe second turbine.
 5. The gas turbine engine as set forth in claim 4,wherein the performance ratio is equal to or greater than 0.8.
 6. Thegas turbine engine as set forth in claim 5, wherein the mid-turbineframe includes a plurality of airfoils in a core airflow path.
 7. Thegas turbine engine as set forth in claim 6, wherein the performanceratio is greater than or equal to 1.0.
 8. The gas turbine engine as setforth in claim 7, wherein the gear system is a planetary gear system. 9.The gas turbine engine as set forth in claim 7, wherein the mid-turbineframe has a second bearing supporting an outer periphery of a secondshaft rotatable with the second turbine.
 10. The gas turbine engine asset forth in claim 9, wherein the gear system is intermediate the fandrive turbine and the first compressor such that the fan and the firstcompressor are rotatable at a common speed.
 11. The gas turbine engineas set forth in claim 5, wherein the performance ratio is less than orequal to 1.075.
 12. The gas turbine engine as set forth in claim 5,wherein the second speed is greater than twice the first speed.
 13. Thegas turbine engine as set forth in claim 1, comprising: a fan driveshaft interconnecting the gear system and the fan; a frame supporting atleast a portion of the fan drive shaft, the frame defining a frametransverse stiffness; a flexible support at least partially supportingthe gear system, the flexible support defining a support transversestiffness with respect to the frame transverse stiffness; and whereinthe support transverse stiffness is less than 50% of the frametransverse stiffness.
 14. The gas turbine engine as set forth in claim1, wherein the gear system is straddle-mounted by bearings.
 15. The gasturbine engine as set forth in claim 14, comprising: a fan drive shaftinterconnecting the gear system and the fan; a frame supporting at leasta portion of the fan drive shaft, the frame defining a frame transversestiffness and a frame lateral stiffness; a flexible support at leastpartially supporting the gear system, the flexible support defining asupport transverse stiffness with respect to the frame transversestiffness and a support lateral stiffness with respect to the framelateral stiffness; wherein the support transverse stiffness is less than80% of the frame transverse stiffness; and wherein the support lateralstiffness is less than 80% of the frame lateral stiffness.
 16. The gasturbine engine as set forth in claim 15, wherein the support transversestiffness is less than 50% of the frame transverse stiffness, and thesupport lateral stiffness is less than 50% of the frame lateralstiffness.
 17. The gas turbine engine as set forth in claim 16, whereinthe support transverse stiffness is less than 20% of the frametransverse stiffness.
 18. The gas turbine engine as set forth in claim17, wherein the support lateral stiffness is less than 20% of the framelateral stiffness.
 19. The gas turbine engine as set forth in claim 18,wherein the second speed is greater than twice the first speed.
 20. Thegas turbine engine as set forth in claim 19, wherein the performanceratio is less than or equal to 1.075.
 21. A gas turbine enginecomprising: a fan having fewer than 26 fan blades; a turbine sectionincluding a fan drive turbine and a second turbine; a gear system with agear reduction, the fan drive turbine driving the fan through the gearsystem; a mid-turbine frame positioned intermediate the fan driveturbine and the second turbine; wherein the mid-turbine frame includes aguide vane positioned intermediate the fan drive turbine and the secondturbine; wherein the fan drive turbine has a first exit area at a firstexit point and is rotatable at a first speed, the second turbine has asecond exit area at a second exit point and is rotatable at a secondspeed, said first and second speeds being redline speeds; and wherein afirst performance quantity is defined as the product of the first speedsquared and the first area, a second performance quantity is defined asthe product of the second speed squared and the second area, and aperformance ratio of the first performance quantity to the secondperformance quantity is between 0.8 and 1.5.
 22. The gas turbine engineas set forth in claim 21, wherein the fan drive turbine and secondturbine are rotatable in opposed directions, and the guide vane is anair turning guide vane.
 23. The gas turbine engine as set forth in claim22, wherein the mid-turbine frame has a first bearing supporting a firstshaft rotatable with the fan drive turbine in an overhung manner. 24.The gas turbine engine as set forth in claim 23, wherein the fan driveturbine is a 3-stage to 6-stage turbine, and the second turbine is a2-stage turbine.
 25. The gas turbine engine as set forth in claim 22,wherein the fan drive turbine is a 3-stage to 6-stage turbine, and thesecond turbine is a 2-stage turbine.
 26. The gas turbine engine as setforth in claim 25, wherein the fan drive turbine is a 3-stage turbine.27. The gas turbine engine as set forth in claim 25, comprising: a fandrive shaft interconnecting the gear system and the fan; a framesupporting at least a portion of the fan drive shaft, the frame defininga frame transverse stiffness; a flexible support at least partiallysupporting the gear system, the flexible support defining a supporttransverse stiffness with respect to the frame transverse stiffness; andwherein the support transverse stiffness is less than 50% of the frametransverse stiffness.
 28. The gas turbine engine as set forth in claim25, wherein the performance ratio is greater than or equal to 1.0. 29.The gas turbine engine as set forth in claim 28, wherein the secondspeed is greater than twice the first speed.